A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine may include, for example, five major sections, a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. The fan section is positioned at the front, or “inlet” section of the engine, and includes a fan that induces air from the surrounding environment into the engine, and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum, and out the exhaust section.
The compressor section raises the pressure of the air it receives from the fan section to a relatively high level. In a multi-spool engine, the compressor section may include two or more compressors. For example, in a triple spool engine, the compressor section may include a high pressure compressor, and an intermediate compressor. The compressed air from the compressor section then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel. The injected fuel is ignited by a burner, which significantly increases the energy of the compressed air.
The high-energy compressed air from the combustor section then flows into and through the turbine section, causing rotationally mounted turbine blades to rotate and generate energy. The air exiting the turbine section is exhausted from the engine via the exhaust section, and the energy remaining in this exhaust air aids the thrust generated by the air flowing through the bypass plenum.
Similar to the compressor section, in a multi-spool (e.g., multi-shaft) engine the turbine section may include a plurality of turbines. For example, in a triple spool engine, the turbine section may include a high pressure turbine, an intermediate pressure turbine, and a low pressure turbine. The energy generated in each of the turbines may be used to power other portions of the engine. For example, the low pressure turbine may be used to power the fan via one spool, the intermediate turbine may be used to power the intermediate pressure turbine via another spool that is concentric to the low pressure turbine spool, and the high pressure turbine may be used to power the high pressure compressor via yet another concentric spool.
Typically, each turbine in a multi-spool gas turbine engine is mounted within a turbine case. The rotationally mounted blades in each turbine extend radially from its associated spool, and are surrounded by, and spaced apart from, an annular seal that is attached to the turbine case. The radial clearance between the turbine blade ends and the annular seal is preferably minimized, in order to reduce leakage of fluid (e.g., air) past the turbine blades. This is because turbine efficiency decreases as fluid leakage past the turbine blades increases.
During a typical gas turbine engine operating cycle, rotational speed and temperature variations within the turbine may result in variations of the radial clearance between the blades and the annular seal. Hence, to avoid contact between the blades and annular seal, the clearance between these components may, under some circumstances, be larger than would otherwise be desirable for certain engine operating conditions. The condition that results in the smallest radial clearance between the blades and annular seal occurs when the gas turbine engine is quickly brought up to full power. This may occur, for example, during an aircraft take-off and climb to cruise altitude. During an aircraft take-off, the turbine blades heat up rapidly and thermally expand. Additionally, the turbine's rotational speed increases, subjecting the blades to centrifugal forces that may cause radial blade growth. At the same time, the annular seal, and the turbine case that supports it, may both heat up rapidly and thermally expand.
After take-off and climb to cruise altitude, when the gas turbine engine is operated in a lower power condition, a temperature equilibrium situation may be reached in the engine. However, the equilibrium temperature reached in various components of the turbines may result in the radial clearance between the turbine blades and annular seal being larger than desirable, resulting in undesirable leakage and concomitantly low engine efficiency.
The rate at which the turbine case and turbine blades thermally expand are preferably matched so that the turbine blade/annular seal radial clearance remains within minimal clearance limits. Attempts have been made to overcome the problem of variation in the radial gap between the sealing member and the blades by providing, in some instances, a sensing and control system. In many of these sensing and control systems, one or more proximity sensors mounted within the turbine case sense the turbine blade/annular seal radial clearance and, in response to the sensed clearance, control the temperature of the turbine case. Many different types of proximity sensors have been, or are being, used to sense turbine blade/annular seal radial clearance. Included among these are capacitance sensors, inductance sensors, optical sensors, acoustic/dynamic response sensors, microwave resonant cavity sensors, and X-ray sensors.
Although safe and generally effective, presently known sensing and control systems suffer certain drawbacks. Namely, the response characteristics, accuracy, and/or reliability of many of the above-mentioned proximity sensor types is not as high as desirable when use in the operating environment of a gas turbine engine. As a result, proximity sensor replacement may be needed more often than is desirable, and turbine efficiency may be decreased, since proximity sensor accuracy may below. Both of these drawbacks can lead to increased system and operational costs.
Hence, there is a need for a turbine blade proximity sensor that addresses one or more of the above-noted drawbacks. Namely, a proximity sensor that exhibits sufficiently high and sufficiently consistent response, accuracy, and reliability in the potentially harsh gas turbine engine environment. The present invention addresses one or more of these needs.